Solar cell arrays are widely used in space as the primary power source for certain spacecraft, such as satellites, due to their reliability and light weight. A typical spacecraft 10 orbiting the earth 12 is illustrated in FIG. 1. Satellite 10 comprises a spacecraft platform 14 from which extends two solar cell wings 16, which form the solar cell array of the satellite. Each solar cell wing 16 comprises a series of panels 18 attached to each other by hinges and attached to the spacecraft by a yoke 20.
Typically, solar cell arrays are stored in a compact manner on the spacecraft prior to launch of the spacecraft and then are deployed once the spacecraft has reached its orbiting altitude. In order to make the solar cell arrays lighter in weight, more compact for storage, and deployable in space, the solar cell panels can be manufactured using a flexible substrate. Various flexible solar cell panels have been used, such as those fabricated from a fiberglass or thin polymeric substrate upon which are bonded numerous thinned crystalline solar cells or from thin metallic foils upon which are fabricated amorphous silicon solar cells.
The general utility of most previously available solar cell arrays has been impaired by the tendency of the solar cell array to vibrate at a low frequency in orbit. The power production of a solar cell panel is directly related to the area of the solar cells that are utilized and, hence, the size of the solar cell panel. Accordingly, it is desirable to make the solar cell panels and solar cell arrays as large as possible to maximize power production. However, as the size of the solar cell panels and the solar cell array increase, the stiffness of the solar cell array decreases and, as a result, the vibration frequency decreases. If the flopping, fluttering and waving of the solar cell array in space is too excessive, the ability of the spacecraft attitude control system to orient the spacecraft is impaired and the spacecraft cannot perform its intended function in orbit. In addition, the efficiency of the solar cell array is reduced and damage to the solar cell array may result.
Previous attempts to achieve adequate stiffness of the solar cell arrays have proven unsatisfactory. On the one hand, power production is compromised if low stiffness is avoided by reducing the size of the solar cell panels and the solar cell array. On the other hand, the use of various stiffeners and rigid structures to stabilize a solar cell array increases the bulk of the solar cell array, making it heavier and more difficult to stow for launch. Moreover, as the complexity of the deployment structures of the large solar cell array increases, the reliability of the structure decreases.
Accordingly, it is desirable to provide a stiff solar cell panel assembly using flexible substrates for deployment from a spacecraft. In addition, it is desirable to provide a deployable solar cell array system that permits use of large solar cell panels. Moreover, it is desirable to provide a solar cell panel assembly that may be compactly stored and may be highly stiff when deployed. It also is desirable to provide a highly reliable deployable solar cell array system. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.